1. Field of the Invention
This invention relates to a reverser flap assembly for a gas turbine nozzle, particularly one with pivoting members.
2. The Prior Art
Reverser flap assemblies for gas turbine nozzles are known in the prior art. See FIG. 1 hereof, and U.S. Pat No. 4,798,328 to Thayer et al (1989) and U.S. Pat. No. 4,828,173 to Guerty (1989), which are incorporated herein by reference.
Per FIG. 1, the prior art nozzle 10 has convergent flap 12 and divergent flap 14 pivotably mounted therein, which can direct axial gas flow through the engine when pivoted to an open position or as here, when convergent flap 12 is pivoted to the closed position, shown in FIG. 1, the engine core gas flow per arrow 15, is directed through the reverser channel 18 and out through cascade vane assembly 20, as shown in FIG. 1. A sliding door 22 slides over the cascade vanes when the reverser flow 15 is not employed, as indicated in FIG. 1.
As shown in FIGS. 2, 3, and 4, the individual vanes 24 and 26 of the prior art vane pack 20, are doubly turned to provide a relatively small opening 28, e.g. as shown in FIGS. 2 and 3. This is because the cascade vanes need to perform the double function of a) reversing the engine jet flow against the direction of travel of an engine (and its aircraft). Also however., b) the vanes 24 of the cascade vane assembly 20 must also turn or splay the so reversed jet flow laterally and outwardly at an angle with the skin of the various aircraft components so as to avoid distorting the air flow thereover and also to avoid re-ingestion of such reverser engine core gas back into the intake of such engine or other engine.
However, this dual turning of cascade vanes 24 and 26, results in a relatively small or constricted air flow path 28 between such vanes, e.g. per FIGS. 2 and 3 hereof, which partially blocks the thrust of the reverser gases (gases 15 of FIG. 1) which requires the construction of larger assemblies of cascade vanes and impairs reverser thrust efficiency. Further, the so-reversed core gas flow, after emerging from the cascade vanes, can, in part, impinge on aft engine control surfaces, and can, in part, travel up the air frame boundary layer, over heating the engine skin, with a possibility of engine inlet re-ingestion of such gas.
U.S. Pat. No. 4,798,328 to Thayer et al (1989) and U.S. Pat. No. 4,828,173 to Guerty (1989), address the reverser problem by employing a plurality of cascade vanes, which pivot at the reverser outlet,adjacent the skin of such engine. There is no directing the reverser gas flow significantly away from the engine skin and its boundary layer, nor is there a teaching of applying lateral vectoring or splay of such reverser effluent away from aircraft control surfaces, which are important considerations in aircraft thrust reversal. That is, the above two patents disclose thrust reverser engines (with reverser cascade vanes) that have similar reverser gas boundary layer and impingement problems.
Accordingly there is a need and market to provide a thrust reverser assembly that largely overcomes the above prior art shortcomings.
There has now been discovered a thrust reverser assembly that provides an effective reverser exhaust path with sufficient pitch and splay angles and with reduced exhaust blockage in a compact package.